1. Technical Field
This invention relates to the internal cooling of gas turbine engine, turbine airfoils and particularly the end portions thereof.
2. Background Art
Modern gas turbine engines operate at temperatures approaching 3000° F. Accordingly, it is a common practice to cool various components employed in such engines with air provided by the engine's compressor. Perhaps the most critical components to cool with compressor air are the first stage turbine blades and vanes which are exposed to products of combustion exiting the engine's combustor.
It is well known to provide such compressor discharge cooling air to first stage turbine blades and vanes by routing such air through passages internally of the airfoil portions thereof. Such passages may be cast into the airfoil portions or drilled into the blades or vanes by mechanical or electrochemical machining processes.
In the case of turbine blades and vanes for large industrial gas turbine engines, it is a common practice to employ shaped tube electrochemical machining to form cooling air passages which extend radially from the inner end of the airfoil to the outer end thereof. For enhanced convective cooling, the cooling air passages often include discontinuities in the walls thereof to enhance the turbulence of the flow of cooling air through the passages by eliminating the boundary layer of airflow along the passage walls. Such discontinuities, often referred to as turbulence promoters or turbulators, may take the form of grooves or ridges in the cooling passage walls.
While such turbulators enhance the convective cooling of the interiors of turbine blades and vanes, they necessarily increase the losses associated with the flow of cooling air through the passages and thus adversely affect the overall efficiency of the engine. Therefore, it has been the conventional wisdom to use such turbulators only where they are most necessary from the standpoint of thermal loading. It is generally accepted in the prior art that the locations where internal cooling of turbine blades and vanes is most critical (where thermal loads are greatest) are those locations in the blade or vane airfoils intermediate the root and tip portions thereof. Accordingly, as a result of qualitative analyses of the operating characteristics of blades and vanes, it has been the practice to provide such turbulators only in the intermediate portions of the internal cooling passages of turbine blades and vanes, the root and tip ends of the passages being smooth to minimize the inefficiencies associated with the creation of turbulent flow therein.
However, inspections of modern industrial gas turbine engines, as part of the routine overhaul and maintenance thereof, has revealed that the blades and vanes of such engines experience significant and often unanticipated thermal stress at the ends thereof as evidenced by, for example, cracking in the blade shrouds, such as, in the fillet where the shroud joins the blade. Several solutions to such thermal stress and damage to the blade have been proposed and typically involve a rather complex distribution of additional cooling passages and chambers in the shroud. While such cooling schemes have met with limited success, they greatly increase the complexity of the internal cooling passage configuration and thus greatly increase the complexity and manufacturing costs of the blade. These increased costs may more than offset the savings in operating costs associated with having smooth bores at the radially inner and outer ends of the airfoil cooling passages.